1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine airfoil with film cooling holes.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
Rotor blades and stator vanes within a turbine section of the gas turbine engine are typically cooled using a combination of convection cooling, impingement cooling and film cooling in order to control a metal temperature of the airfoil and prevent hot spots from occurring that can lead to erosion damage and therefore a short part life. This is especially critical in the industrial engines, since these engines must operate continuously for long periods of time.
Film cooling is used to discharge a blanket of film cooling air over the external surface of the airfoil and prevent the hot gas stream from contacting the airfoil external surface. Film cooling holes are mainly used on the airfoil leading edge region surface which is the surface of the airfoil exposed to the highest gas stream temperature. Large length to diameter film cooling holes are used in the leading edge region to provide both internal convection cooling to the airfoil wall and external film cooling for the external surface. For a laser or EDM (electric discharge machining) film cooling hole, a typical length to diameter ratio is less than 12 and the film cooling hole angle is usually no less than 20 degrees relative to the airfoil leading edge surface. FIG. 1 show a prior art film cooling hole with a large L/D ratio and is a straight hole with a constant diameter from an inlet end to an outlet end that provides no diffusion of the film cooling air prior to discharge.
FIG. 2 shows a prior art film cooling air hole with a constant diameter metering inlet section 11 and a diffusion section 12 that opens onto the airfoil surface. this film cooling hole is angled at 25 degrees to the airfoil surface with a 10 degree expansion on the downstream wall of the diffusion section 12. Both the film cooling holes in FIGS. 1 and 2 have an L/D ratio of around 14 and both film cooling holes have hole angles and L/D ratios that exceed current manufacturing capability. Because of the diffusion section in the FIG. 2 film cooling hole, a large film hole breakout is formed on the airfoil surface. U.S. Pat. No. 6,869,268 issued to Liang on Mar. 22, 2005 and entitled COMBUSTION TURBINE WITH AIRFOIL HAVING ENHANCED LEADING EDGE DIFFUSION HOLES AND RELATED METHODS discloses the FIG. 2 film cooling hole.
A further improvement of the film cooling holes is shown in FIGS. 3 and 4 in which the constant diameter metering section 21 discharges into a first diffusion section 22 and then a second diffusion section 23 that opens onto the airfoil surface. U.S. Pat. No. 4,653,983 issued to Vehr on Mar. 31, 1987 and entitled CROSS-FLOW FILM COOLING PASSAGES and U.S. Pat. No. 5,382,133 issued to Moore et al on Jan. 17, 1995 and entitled HIGH COVERAGE SHAPED DIFFUSER FILM HOLE FOR THIN WALLS discloses these types of double diffusion film cooling holes.
U.S. Pat. No. 4,684,323 issued to Field on Aug. 4, 1987 and entitled FILM COOLING PASSAGES WITH CURVED CORNERS and U.S. Pat. No. 6,183,199 issued to Beeck et al on Feb. 6, 2001 and entitled COOLING-AIR BORE discloses three dimension holes in an axial or small compound angle and a variety of expansion shapes that further enhances the film cooling capability.
A further improvement over the three-dimensional diffusion holes is disclosed in U.S. Pat. No. 6,918,742 issued to Liang on Jul. 19, 2005 and entitled COMBUSTION TURBINE WITH AIRFOIL HAVING MULTI-SECTION DIFFUSION COOLING HOLES AND METHODS OF MAKING SAME which discloses a multiple diffusion compounded film cooling holes having a constant diameter metering inlet section to provide cooling flow metering capability followed by a 3 to 5 degree expansion in the radial outward direction and a combination of 3 to 5 degree followed by a 10 degree multi-expansion in the downstream and radial inboard directions. There is no expansion for the film hole on the upstream side wall where the film cooling hole is in contact with the hot gas stream.